Airfoil having impingement leading edge

ABSTRACT

Airfoils for a gas turbine engines and gas turbine engines are described. The airfoils include an airfoil body extending in a radial direction from a first end to a second end, and axially from a leading edge to a trailing edge. A radially extending leading edge channel is formed in the leading edge of the airfoil body, having first and second channel walls that join at a channel base. A first leading edge impingement cavity is located within the airfoil body proximate the leading edge and is defined, in part, by the first channel wall. A leading edge feed cavity is arranged aft of the first leading edge impingement cavity to supply air into the first leading edge impingement cavity. A first leading edge impingement hole is formed in the first channel wall and angled toward a portion of the second channel wall.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a continuation application of U.S. application Ser.No. 15/791,464 filed Oct. 24, 2017, the contents of which areincorporated by reference herein in their entirety.

BACKGROUND

Illustrative embodiments pertain to the art of turbomachinery, andspecifically to turbine rotor components.

Gas turbine engines are rotary-type combustion turbine engines builtaround a power core made up of a compressor, combustor and turbine,arranged in flow series with an upstream inlet and downstream exhaust.The compressor compresses air from the inlet, which is mixed with fuelin the combustor and ignited to generate hot combustion gas. The turbineextracts energy from the expanding combustion gas, and drives thecompressor via a common shaft. Energy is delivered in the form ofrotational energy in the shaft, reactive thrust from the exhaust, orboth.

The individual compressor and turbine sections in each spool aresubdivided into a number of stages, which are formed of alternating rowsof rotor blade and stator vane airfoils. The airfoils are shaped toturn, accelerate and compress the working fluid flow, or to generatelift for conversion to rotational energy in the turbine.

Airfoils may incorporate various cooling cavities located adjacentexternal side walls. Such cooling cavities are subject to both hotmaterial walls (exterior or external) and cold material walls (interioror internal). Although such cavities are designed for cooling portionsof airfoil bodies, various cooling flow characteristics can cause hotsections where cooling may not be sufficient. Accordingly, improvedmeans for providing cooling for an airfoil may be desirable.

BRIEF DESCRIPTION

According to some embodiments, airfoils for gas turbine engines areprovided. The airfoils have an airfoil body extending in a radialdirection from a first end to a second end, and extending axially from aleading edge to a trailing edge, a leading edge channel formed in theleading edge of the airfoil body, the leading edge channel having afirst channel wall and a second channel wall that join at a channel baseto define the leading edge channel, the leading edge channel extendingin a radial direction along the leading edge of the airfoil body, afirst leading edge impingement cavity located within the airfoil bodyproximate the leading edge, wherein the first channel wall forms aportion of the airfoil body defining the first leading edge impingementcavity, a second leading edge impingement cavity located within theairfoil body proximate the leading edge, wherein the second channel wallforms a portion of the airfoil body defining the second leading edgeimpingement cavity, and a first leading edge impingement hole formed inthe first channel wall and angled such that air flowing from the firstleading edge impingement cavity and through the first leading edgeimpingement hole impinges upon a portion of the second channel wall.

In addition to one or more of the features described herein, or as analternative, further embodiments of the airfoils may include that thefirst end is a root of the airfoil body and the second end is a tip ofthe airfoil, wherein the leading edge channel extends less than a fulllength of a distance between the root and the tip along the leading edgeof the airfoil body.

In addition to one or more of the features described herein, or as analternative, further embodiments of the airfoils may include a secondleading edge impingement hole formed in the second channel wall andangled such that air flowing from the second leading edge impingementcavity and through the second leading edge impingement hole impingesupon a portion of the first channel wall.

In addition to one or more of the features described herein, or as analternative, further embodiments of the airfoils may include that theleading edge channel comprises a plurality of first leading edgeimpingement holes formed in the first channel wall, wherein theplurality of first leading edge impingement holes extend in an arrayradially along the first channel wall, and wherein the leading edgechannel comprises a plurality of second leading edge impingement holesformed in the second channel wall, wherein the plurality of secondleading edge impingement holes extend in an array radially along thesecond channel wall.

In addition to one or more of the features described herein, or as analternative, further embodiments of the airfoils may include a leadingedge feed cavity arranged aft of the first and second leading edgeimpingement cavities and arranged to supply air into at least one of thefirst leading edge impingement cavity and the second leading edgeimpingement cavity.

In addition to one or more of the features described herein, or as analternative, further embodiments of the airfoils may include that theleading edge feed cavity supplies air into both of the first leadingedge impingement cavity and the second leading edge impingement cavity.

In addition to one or more of the features described herein, or as analternative, further embodiments of the airfoils may include that theleading edge channel had a depth in an axial direction that is at leasttwice a diameter of the first leading edge impingement hole.

In addition to one or more of the features described herein, or as analternative, further embodiments of the airfoils may include that theleading edge channel comprises a plurality of first leading edgeimpingement holes formed in the first channel wall, wherein theplurality of first leading edge impingement holes extend in an arrayradially along the first channel wall.

In addition to one or more of the features described herein, or as analternative, further embodiments of the airfoils may include at leastone additional channel formed on an exterior surface of the airfoil andextending radially.

In addition to one or more of the features described herein, or as analternative, further embodiments of the airfoils may include that theairfoil body forms a blade of a turbine section of the gas turbineengine.

According to some embodiments, gas turbine engines are provided. The gasturbine engines include an airfoil having an airfoil body extending in aradial direction from a first end to a second end, and extending axiallyfrom a leading edge to a trailing edge, a leading edge channel formed inthe leading edge of the airfoil body, the leading edge channel having afirst channel wall and a second channel wall that join at a channel baseto define the leading edge channel, the leading edge channel extendingin a radial direction along the leading edge of the airfoil body, afirst leading edge impingement cavity located within the airfoil bodyproximate the leading edge, wherein the first channel wall forms aportion of the airfoil body defining the first leading edge impingementcavity, a second leading edge impingement cavity located within theairfoil body proximate the leading edge, wherein the second channel wallforms a portion of the airfoil body defining the second leading edgeimpingement cavity, and a first leading edge impingement hole formed inthe first channel wall and angled such that air flowing from the firstleading edge impingement cavity and through the first leading edgeimpingement hole impinges upon a portion of the second channel wall.

In addition to one or more of the features described herein, or as analternative, further embodiments of the gas turbine engines may includethat the first end is a root of the airfoil body and the second end is atip of the airfoil, wherein the leading edge channel extends less than afull length of a distance between the root and the tip along the leadingedge of the airfoil body.

In addition to one or more of the features described herein, or as analternative, further embodiments of the gas turbine engines may includea second leading edge impingement hole formed in the second channel walland angled such that air flowing from the second leading edgeimpingement cavity and through the second leading edge impingement holeimpinges upon a portion of the first channel wall.

In addition to one or more of the features described herein, or as analternative, further embodiments of the gas turbine engines may includethat the leading edge channel comprises a plurality of first leadingedge impingement holes formed in the first channel wall, wherein theplurality of first leading edge impingement holes extend in an arrayradially along the first channel wall, and wherein the leading edgechannel comprises a plurality of second leading edge impingement holesformed in the second channel wall, wherein the plurality of secondleading edge impingement holes extend in an array radially along thesecond channel wall.

In addition to one or more of the features described herein, or as analternative, further embodiments of the gas turbine engines may includea leading edge feed cavity arranged aft of the first and second leadingedge impingement cavities and arranged to supply air into at least oneof the first leading edge impingement cavity and the second leading edgeimpingement cavity.

In addition to one or more of the features described herein, or as analternative, further embodiments of the gas turbine engines may includethat the leading edge feed cavity supplies air into both of the firstleading edge impingement cavity and the second leading edge impingementcavity.

In addition to one or more of the features described herein, or as analternative, further embodiments of the gas turbine engines may includethat the leading edge channel had a depth in an axial direction that isat least twice a diameter of the first leading edge impingement hole.

In addition to one or more of the features described herein, or as analternative, further embodiments of the gas turbine engines may includethat the leading edge channel comprises a plurality of first leadingedge impingement holes formed in the first channel wall, wherein theplurality of first leading edge impingement holes extend in an arrayradially along the first channel wall.

In addition to one or more of the features described herein, or as analternative, further embodiments of the gas turbine engines may includeat least one additional channel formed on an exterior surface of theairfoil and extending radially.

In addition to one or more of the features described herein, or as analternative, further embodiments of the gas turbine engines may includethat the airfoil is a blade of a turbine section of the gas turbineengine.

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be understood, however, the following descriptionand drawings are intended to be illustrative and explanatory in natureand non-limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

The following descriptions should not be considered limiting in any way.With reference to the accompanying drawings, like elements are numberedalike: The subject matter is particularly pointed out and distinctlyclaimed at the conclusion of the specification. The foregoing and otherfeatures, and advantages of the present disclosure are apparent from thefollowing detailed description taken in conjunction with theaccompanying drawings in which like elements may be numbered alike and:

FIG. 1 is a schematic cross-sectional illustration of a gas turbineengine;

FIG. 2 is a schematic illustration of a portion of a turbine section ofthe gas turbine engine of FIG. 1;

FIG. 3A is a schematic illustration of an airfoil in accordance with anembodiment of the present disclosure; and

FIG. 3B is an enlarged illustration of a leading edge of the airfoilshown in FIG. 3A.

DETAILED DESCRIPTION

Detailed descriptions of one or more embodiments of the disclosedapparatus and/or methods are presented herein by way of exemplificationand not limitation with reference to the Figures.

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct, while the compressor section 24 drives air along a coreflow path C for compression and communication into the combustor section26 then expansion through the turbine section 28. Although depicted as atwo-spool turbofan gas turbine engine in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with two-spool turbofans as the teachings may beapplied to other types of turbine engines including three-spoolarchitectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through aspeed change mechanism, which in exemplary gas turbine engine 20 isillustrated as a geared architecture 48 to drive the fan 42 at a lowerspeed than the low speed spool 30. The high speed spool 32 includes anouter shaft 50 that interconnects a high pressure compressor 52 and highpressure turbine 54. A combustor 56 is arranged in exemplary gas turbine20 between the high pressure compressor 52 and the high pressure turbine54. An engine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The engine staticstructure 36 further supports bearing systems 38 in the turbine section28. The inner shaft 40 and the outer shaft 50 are concentric and rotatevia bearing systems 38 about the engine central longitudinal axis Awhich is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion. It will be appreciated that each of the positions of the fansection 22, compressor section 24, combustor section 26, turbine section28, and fan drive gear system 48 may be varied. For example, gear system48 may be located aft of combustor section 26 or even aft of turbinesection 28, and fan section 22 may be positioned forward or aft of thelocation of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present disclosure isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and35,000 ft (10,688 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(514.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5 m/sec).

Although the gas turbine engine 20 is depicted as a turbofan, it shouldbe understood that the concepts described herein are not limited to usewith the described configuration, as the teachings may be applied toother types of engines such as, but not limited to, turbojets,turboshafts, and three-spool (plus fan) turbofans wherein anintermediate spool includes an intermediate pressure compressor (“IPC”)between a low pressure compressor (“LPC”) and a high pressure compressor(“HPC”), and an intermediate pressure turbine (“IPT”) between the highpressure turbine (“HPT”) and the low pressure turbine (“LPT”).

FIG. 2 is a schematic view of a portion of the turbine section 28 thatmay employ various embodiments disclosed herein. Turbine section 28includes a plurality of airfoils 60, 62 including, for example, one ormore blades and vanes. The airfoils 60, 62 may be hollow bodies withinternal cavities defining a number of channels or cores, hereinafterairfoil cores, formed therein and extending from an inner diameter 66 toan outer diameter 68, or vice-versa. The airfoil cores may be separatedby partitions within the airfoils 60, 62 that may extend either from theinner diameter 66 or the outer diameter 68 of the airfoil 60, 62. Thepartitions may extend for a portion of the length of the airfoil 60, 62,but may stop or end prior to forming a complete wall within the airfoil60, 62. Thus, each of the airfoil cores may be fluidly connected andform a fluid path within the respective airfoil 60, 62. The airfoils 60,62 may include platforms 70 located proximal to the inner diameter 66thereof. Located below the platforms 70 (e.g., radially inward withrespect to the engine axis) may be airflow ports and/or bleed orificesthat enable air to bleed from the internal cavities of the airfoils 60,62. A root of the airfoil may connect to or be part of the platform 70.

The turbine section 28 is housed within a case 80, which may havemultiple parts (e.g., turbine case, diffuser case, etc.). In variouslocations, components, such as seals, may be positioned between airfoils60, 62 and the case 80. For example, as shown in FIG. 2, blade outer airseals 82 (hereafter “BOAS”) are located radially outward from the blade60. As will be appreciated by those of skill in the art, the BOAS 82 mayinclude BOAS supports that are configured to fixedly connect or attachthe BOAS 82 to the case 80 (e.g., the BOAS supports may be locatedbetween the BOAS 82 and the case 80). As shown in FIG. 2, the case 80includes a plurality of case hooks 84 that engage with BOAS hooks 86 tosecure the BOAS 82 between the case 80 and a tip of the airfoil 60.

As shown and labeled in FIG. 2, a radial direction is upward on the page(e.g., radial with respect to an engine axis) and an axial direction isto the right on the page (e.g., along an engine axis). Thus, radialcooling flows will travel up or down on the page and axial flows willtravel left-to-right (or vice versa).

Turning to FIGS. 3A-3B, schematic illustrations of an airfoil 300 inaccordance with an embodiment of the present disclosure are shown, withFIG. 3B being an enlarged illustration of a leading edge of the airfoil300. The airfoil 300 is defined by an airfoil body 300 a having aleading edge 302 and extends aftward to a trailing edge 304. The airfoil300, as shown, has a first leading edge impingement cavity 306 and asecond leading edge impingement cavity 308. Aft of the leading edgeimpingement cavities 306, 308 is a leading edge feed cavity 310. Theairfoil 300 includes additional cooling cavities 312, as shown.

The leading edge impingement cavities 306, 308 are fed from the leadingedge feed cavity 310 by one or more respective impingement holes thatfluidly connect the leading edge feed cavity 310 leading edgeimpingement cavities 306, 308. For example, as shown, a firstimpingement hole 314 enables air from the leading edge feed cavity 310to impinge into the first leading edge impingement cavity 306 and asecond impingement hole 316 enables air from the leading edge feedcavity 310 to impinge into the second leading edge impingement cavity308. The impinging air from the leading edge feed cavity 310 willprovide cooling to the leading edge 302 of the airfoil 300.

As shown, the airfoil 300 includes a stagnation divot, hollow, trench,or channel (hereinafter “leading edge channel 318”) which is positionedat the tip or front of the leading edge 302 of the airfoil 300. Theleading edge channel 318, in some embodiments, extends from a root to atip (e.g., along a radial length) of the airfoil 300, as will beappreciated by those of skill in the art. In other embodiments, theleading edge channel can extend over a partial extent of the radiallength, e.g., less than a full length of a distance between the firstend (e.g., root) and the second (e.g., tip) along the leading edge ofthe airfoil body. That is, in some embodiments, the leading edge channelmay not extend along the entire radial length of the leading edge. Forexample, in one non-limiting embodiment, the leading edge channel mayextend along about half of the radial length of the airfoil leadingedge, with no channel present on the other half of the leading edge. Inone such embodiment, the leading edge channel can extend from a tip ofthe airfoil in a radially downward direction and stop approximately atthe halfway point. In other embodiments, the leading edge channel can beless-than-full-radial-distance in length, but positioned such thatneither end of the channel is located at the root or tip (e.g.,positioned at a mid-point or partial mid-point along the radial lengthof the leading edge of the airfoil). Further, in some embodiments,multiple less-than-full length leading edge channels can be formed on aleading edge of the airfoil, without departing from the scope of thepresent disclosure.

A portion of the air from the first and second leading edge impingementcavities 306, 308 will exit the respective first and second leading edgeimpingement cavities 306, 308 through leading edge impingement holes320, 322. A first leading edge impingement hole 320 is formed in a firstchannel wall 324 that partially defines the first leading edgeimpingement cavity 306. A second leading edge impingement hole 322 isformed in a second channel wall 326 that partially defines the secondleading edge impingement cavity 308. Although a single leading edgeimpingement hole is shown in the illustration of FIGS. 3A-3B, those ofskill in the art will appreciate that an array of impingement holes canbe formed extending along a radial extent of the airfoil 300 along thechannel walls 324, 326 of the leading edge channel. The first channelwall 324 and the second channel wall 326 form and define the leadingedge channel 318. A channel base 328 is formed where the first andsecond channel walls 324, 326 meet at the base of the leading edgechannel 318.

Air from the first and second leading edge impingement cavities 306, 308will flow into and/or impinge into the leading edge channel 318 alongthe leading edge 302. Once in the leading edge channel 318, the coolingair diffuses into cooling air already in the leading edge channel 318and distributes spanwise along the leading edge channel 318. One of theadvantages of distributing cooling air within the leading edge channel318 is that the pressure difference problems characteristic ofconventional cooling orifices are minimized. For example, the differencein pressure across a cooling orifice is a function of a local internalcavity pressure and a local gaspath gas pressure adjacent the orifice.Both of these pressures vary as a function of time. If the gaspath gaspressure is high and the internal cavity pressure is low adjacent aparticular cooling orifice in a conventional scheme, undesirable hotcore gas in-flow can occur (e.g., into one of the leading edgeimpingement or other cooling cavities). Embodiments provided herein canminimize the opportunity for the undesirable in-flow due to impingementair being distributed within the leading edge channel 318, therebydecreasing the opportunity for any low pressure zones to occur.Likewise, the distribution of cooling air within the leading edgechannel 318 also avoids cooling air pressure spikes which, in aconventional scheme, can jet cooling air into the gaspath gas ratherthan add it to a film of cooling air downstream along exterior surfacesof the airfoil 300.

Additionally, air impinging from the leading edge impingement holes 320,322 can provide impingement cooling to the opposing channel walls 324,326. For example, as shown in FIG. 3B, air from the first leading edgeimpingement cavity 306 will flow through the first leading edgeimpingement hole 320 and impinge upon the second channel wall 326.Similarly, air from the second leading edge impingement cavity 308 willflow through the second leading edge impingement hole 322 and impingeupon the first channel wall 324. That is, the first leading edgeimpingement hole 320 is angled such that air passing through the firstleading edge impingement hole 320 is directed such that air will impingeupon the material of the second channel wall 326 and the second leadingedge impingement hole 322 is angled such that air passing through thesecond leading edge impingement hole 322 is directed such that air willimpinge upon the material of the first channel wall 324.

In some embodiments, the leading edge channel 318 has a depth 330 thatis at least twice the diameter 332 of the leading edge impingement holes320, 322, as schematically shown. This dimension enables formation ofthe leading edge impingement holes 320, 322 within the channel walls324, 326. The depth 330 is measured from the leading edge 302 to thechannel base 328 of the leading edge channel 318.

Although shown herein with two leading edge impingement holes (one foreach leading edge impingement cavity), those of skill in the art willappreciate that other arrangements are possible without departing fromthe scope of the present disclosure. For example, in some embodiments,one of the leading edge impingement holes can be omitted such that onlyone leading edge impingement hole is provided to supply air into theleading edge channel. Further, in some embodiment, multiple channels canbe formed along the leading edge or other surfaces of the airfoil. Insome such embodiments, each of the channels can be arranged with one ormore impingement and/or feed holes to supply air into the channel, aswill be appreciated by those of skill in the art. For example, in onenon-limiting embodiment, a second channel can be formed on the exteriorsurface of the airfoil 300 adjacent the first leading edge impingementcavity 306 and the leading edge feed cavity 310, with the second channelbeing sourced from the first leading edge impingement cavity 306 and theleading edge feed cavity 310.

Embodiments provided herein are directed to airfoils having a leadingedge channel that is supplied with impingement air from multipledifferent leading edge impingement cavities. Advantageously, sucharrangement can enable the leading edge channel to be supplied withimpingement air from any one (or more) of the leading edge impingementcavities, thus ensuring constant impinging air within the leading edgechannel. Further, advantageously, embodiment provided herein aredirected to angled impingement holes within the leading edge channelsuch that a portion of the impinging air that flows into the leadingedge channel impinges upon and cools the material of the opposingchannel wall.

As used herein, the term “about” is intended to include the degree oferror associated with measurement of the particular quantity based uponthe equipment available at the time of filing the application. Forexample, “about” may include a range of ±8%, or 5%, or 2% of a givenvalue or other percentage change as will be appreciated by those ofskill in the art for the particular measurement and/or dimensionsreferred to herein.

The terminology used herein is for the purpose of describing particularembodiments only and is not intended to be limiting of the presentdisclosure. As used herein, the singular forms “a,” “an,” and “the” areintended to include the plural forms as well, unless the context clearlyindicates otherwise. It will be further understood that the terms“comprises” and/or “comprising,” when used in this specification,specify the presence of stated features, integers, steps, operations,elements, and/or components, but do not preclude the presence oraddition of one or more other features, integers, steps, operations,element components, and/or groups thereof. It should be appreciated thatrelative positional terms such as “forward,” “aft,” “upper,” “lower,”“above,” “below,” “radial,” “axial,” “circumferential,” and the like arewith reference to normal operational attitude and should not beconsidered otherwise limiting.

While the present disclosure has been described with reference to anillustrative embodiment or embodiments, it will be understood by thoseskilled in the art that various changes may be made and equivalents maybe substituted for elements thereof without departing from the scope ofthe present disclosure. In addition, many modifications may be made toadapt a particular situation or material to the teachings of the presentdisclosure without departing from the essential scope thereof.Therefore, it is intended that the present disclosure not be limited tothe particular embodiment disclosed as the best mode contemplated forcarrying out this present disclosure, but that the present disclosurewill include all embodiments falling within the scope of the claims.

What is claimed is:
 1. An airfoil for a gas turbine engine, the airfoilcomprising: an airfoil body extending in a radial direction from a firstend to a second end, and extending axially from a leading edge to atrailing edge; a leading edge channel formed in the leading edge of theairfoil body, the leading edge channel having a first channel wall and asecond channel wall that join at a channel base to define the leadingedge channel, the leading edge channel extending in a radial directionalong the leading edge of the airfoil body; a first leading edgeimpingement cavity located within the airfoil body proximate the leadingedge, wherein the first channel wall forms a portion of the airfoil bodydefining the first leading edge impingement cavity; a leading edge feedcavity arranged aft of the first leading edge impingement cavity andarranged to supply air into the first leading edge impingement cavity;and a first leading edge impingement hole formed in the first channelwall and angled such that air flowing from the first leading edgeimpingement cavity and through the first leading edge impingement holeimpinges upon a portion of the second channel wall.
 2. The airfoil ofclaim 1, wherein the first end is a root of the airfoil body and thesecond end is a tip of the airfoil, wherein the leading edge channelextends less than a full length of a distance between the root and thetip along the leading edge of the airfoil body.
 3. The airfoil of claim1, further comprising a second leading edge impingement cavity locatedwithin the airfoil body proximate the leading edge, wherein the secondchannel wall forms a portion of the airfoil body defining the secondleading edge impingement cavity.
 4. The airfoil of claim 3, furthercomprising a second leading edge impingement hole formed in the secondchannel wall and angled such that air flowing from the second leadingedge impingement cavity and through the second leading edge impingementhole impinges upon a portion of the first channel wall.
 5. The airfoilof claim 4, wherein the leading edge channel comprises a plurality offirst leading edge impingement holes formed in the first channel wall,wherein the plurality of first leading edge impingement holes extend inan array radially along the first channel wall, and wherein the leadingedge channel comprises a plurality of second leading edge impingementholes formed in the second channel wall, wherein the plurality of secondleading edge impingement holes extend in an array radially along thesecond channel wall.
 6. The airfoil of claim 3, wherein the leading edgefeed cavity is arranged aft of the second leading edge impingementcavity and arranged to supply air into the second leading edgeimpingement cavity.
 7. The airfoil of claim 1, wherein the leading edgechannel has a depth in an axial direction that is at least twice adiameter of the first leading edge impingement hole.
 8. The airfoil ofclaim 1, wherein the leading edge channel comprises a plurality of firstleading edge impingement holes formed in the first channel wall, whereinthe plurality of first leading edge impingement holes extend in an arrayradially along the first channel wall.
 9. The airfoil of claim 1,further comprising at least one additional channel formed on an exteriorsurface of the airfoil and extending radially.
 10. The airfoil of claim1, wherein the airfoil body forms a blade of a turbine section of thegas turbine engine.
 11. A gas turbine engine comprising: an airfoilhaving an airfoil body extending in a radial direction from a first endto a second end, and extending axially from a leading edge to a trailingedge; a leading edge channel formed in the leading edge of the airfoilbody, the leading edge channel having a first channel wall and a secondchannel wall that join at a channel base to define the leading edgechannel, the leading edge channel extending in a radial direction alongthe leading edge of the airfoil body; a first leading edge impingementcavity located within the airfoil body proximate the leading edge,wherein the first channel wall forms a portion of the airfoil bodydefining the first leading edge impingement cavity; a leading edge feedcavity arranged aft of the first leading edge impingement cavity andarranged to supply air into the first leading edge impingement cavity;and a first leading edge impingement hole formed in the first channelwall and angled such that air flowing from the first leading edgeimpingement cavity and through the first leading edge impingement holeimpinges upon a portion of the second channel wall.
 12. The gas turbineengine of claim 11, wherein the first end is a root of the airfoil bodyand the second end is a tip of the airfoil, wherein the leading edgechannel extends less than a full length of a distance between the rootand the tip along the leading edge of the airfoil body.
 13. The gasturbine engine of claim 11, further comprising a second leading edgeimpingement cavity located within the airfoil body proximate the leadingedge, wherein the second channel wall forms a portion of the airfoilbody defining the second leading edge impingement cavity.
 14. The gasturbine engine of claim 13, further comprising a second leading edgeimpingement hole formed in the second channel wall and angled such thatair flowing from the second leading edge impingement cavity and throughthe second leading edge impingement hole impinges upon a portion of thefirst channel wall.
 15. The gas turbine engine of claim 14, wherein theleading edge channel comprises a plurality of first leading edgeimpingement holes formed in the first channel wall, wherein theplurality of first leading edge impingement holes extend in an arrayradially along the first channel wall, and wherein the leading edgechannel comprises a plurality of second leading edge impingement holesformed in the second channel wall, wherein the plurality of secondleading edge impingement holes extend in an array radially along thesecond channel wall.
 16. The gas turbine engine of claim 13, wherein theleading edge feed cavity is arranged aft of the second leading edgeimpingement cavity and arranged to supply air into the second leadingedge impingement cavity.
 17. The gas turbine engine of claim 11, whereinthe leading edge channel has a depth in an axial direction that is atleast twice a diameter of the first leading edge impingement hole. 18.The gas turbine engine of claim 11, wherein the leading edge channelcomprises a plurality of first leading edge impingement holes formed inthe first channel wall, wherein the plurality of first leading edgeimpingement holes extend in an array radially along the first channelwall.
 19. The gas turbine engine of claim 11, further comprising atleast one additional channel formed on an exterior surface of theairfoil and extending radially.
 20. The gas turbine engine of claim 11,wherein the airfoil is a blade of a turbine section of the gas turbineengine.